The invention relates to propellant designs in Micro Pulsed Plasma Thrusters, and particularly to techniques for bundling propellant modules and for using a two-stage discharge process to increase MicroPPT propellant throughput, and decrease the output voltage required from the power-processing unit.
Micro Pulsed Plasma Thrusters (MicroPPTs) are small thrusters used for propulsion and attitude control on spacecraft. The MicroPPTs are designed to provide spacecraft propulsion at low power levels ( less than 20 W), in small reproducible impulse bits of thrust, and be contained in a very small lightweight package ( less than 200 grams). For 100-kg class microsatellites, MicroPPTs can provide propulsive attitude control and at least partial stationkeeping propulsion. For 25-kg class or smaller satellites, MicroPPTs can provide all primary propulsion, as well as stationkeeping and attitude control.
While MicroPPTs retain some design similarity with standard PPTs, they are fundamentally different in critical areas that enable the required reductions in mass, size, and power. For example, the LES-8/9 PPT, designed and flight tested in the 1970s and shown schematically in FIG. 1 and in the photograph of FIG. 2, a DCxe2x80x94DC converter charges an integrated capacitor from a 28-V spacecraft bus to 1500 V. Then, a second DCxe2x80x94DC converter supplies 600 V to a smaller capacitor in the trigger circuit.
Modern flight units operate in a similar regime. The PPT discharge is initiated by a TTL pulse applied to a semiconductor switch in a trigger circuit. The trigger discharge fires a sparkplug embedded in the cathode, providing enough surface ionization or seed plasma to initiate a main discharge across a Teflon(trademark) propellant face. The solid propellant is converted to vapor and partially ionized by the electric discharge. Acceleration is accomplished by a combination of thermal and electromagnetic forces to create usable thrust. As the propellant is consumed, over some 17 million discharges, a negator spring passively feeds the 25-cm-long propellant bar forward between the electrodes.
Although such systems have been well-tested, the required integration of two separate discharge circuits and two separate DCxe2x80x94DC converters dramatically increases the size of the discharge capacitor and the complexity of the PPT propulsion system. MicroPPTs reduce the need for these duplicative and complicated triggering electronics schemes.
Accordingly, the primary difference between the standard PPT and the MicroPPT lies in the electronics. By reducing the size of the thruster, it becomes possible to initiate the discharge by simple over-voltage breakdown at the propellant face, rather than requiring a separate spark trigger. This eliminates one of the two separate discharge circuits and DCxe2x80x94DC converters required in a conventional PPT, and dramatically reduces the size of the discharge capacitor.
However, although the electronics for these MicroPPTs are much improved, the MicroPPT propellant throughput is fundamentally limited. Adding propellant requires the use of a longer propellant rod or a larger diameter propellant rod. A longer rod becomes impractical for spacecraft integration after approximately 12 inches from a geometric packing perspective. Furthermore, for designs where the propellant recesses back into the electrode shell over time, viscous losses with the wall may decrease thrust. In addition, larger diameter propellants are problematic for the DCxe2x80x94DC converter since higher voltages are needed to initiate the surface discharge.
For example, Laboratory tests have defined two critical design criteria for MicroPPT propellants. First, the voltage required to initiate the surface breakdown increases with propellant diameter since the path length between the inner and outer electrodes increases. Second, the average energy per discharge pulse must increase with propellant face area in order to facilitate complete decomposition of the propellant surface. Insufficient energy leads to charring at the propellant face, which dramatically increases the voltage required to initiate the discharge. This voltage increase is considered a failure mechanism for the MicroPPT.
Accordingly, a need exists for an improved propellant design with large cross-sectional area that also has a lower surface breakdown voltage.
The present invention relates to propellant designs in Micro Pulsed Plasma Thrusters, and particularly to techniques for creating propellant modules capable of increasing MicroPPT propellant throughput, and decreasing the output voltage required from the power-processing unit.
In one embodiment, several modules of relatively smaller diameter are fabricated into a bundle. Such devices are referred to as xe2x80x9cBundledxe2x80x9d propellant designs. In such a xe2x80x9cbundledxe2x80x9d embodiment the MicroPPT could be designed to use one propellant rod until consumed, and then passively switch to the next one.
In another embodiment, a single module of relatively large cross-section is utilized and ignited using a two-stage ignition process. Such devices are referred to as xe2x80x9cTwo-Stagexe2x80x9d propellant designs.
In yet another embodiment, the invention is directed to thrusters utilizing such propellant modules. The propellant modules of the current invention may be utilized in either MicroPPT or Pulsed Plasma Thruster (PPT) designs. In addition, the propellant modules of the current invention may be utilized in any of the MicroPPT designs listed above and only differ in how the high voltage pulse is electrically generated and then applied to the propellant.